The present invention refers to a leading edge of supporting surfaces of aircraft, such as wings, stabilisers and control surfaces, with a ribless structure and a specific frontal area against impacts.
The leading edge may be of a fixed, detachable, articulated or mixed type, and may include detachable panels or access covers.
The manufacture of leading edges in accordance with the present invention is based on theoretical concepts which have been supported by trials and practical developments and which finally have allowed manufacture and certification on commercial aircraft.
The definition of the leading edge is divided conceptually into two parts:
One part is basically structural, called the primary structure, which occupies most of the leading edge and is made up of almost all the upper and lower covering panels of said leading edge and two auxiliary spars, at least one of which is diagonal and makes up the basic conceptual element which allows the classic multi-rib structure to be eliminated from leading edges.
The second part, called frontal area, has a more local character but with a very specific function. In effect, its geometry serves to close the leading edge at the front, although its main object is to withstand erosion and the different types of impact which, as is known, are the most typical problems of this area.
The present invention is basically oriented towards all types of leading edge structures as elements belonging to supporting surfaces of aircraft, and it has a valid application in leading edges with a conventional metallic structure. However, the invention has been developed mainly with regard to the partial or total use of composite materials, especially carbon fibre.
The materials which can be used in the two conceptual parts mentioned above (primary structure and frontal area) may have a wide spectrum and a high degree of combination with each other. The most important, according to their configuration, are as follows:
Upper and lower covering panels
Metallic materials (plate with stiffeners or mechanised panels).
Composite materials: (xe2x80x9csandwichxe2x80x9d type or solid laminate with stiffeners, obtained in their different forms: tape, fabric, RTM (Resin Transfer Moulding), RFI (Resin Film Infusion) or thermoplastics).
Diagonal and frontal spars
Metallic material (plate rigidified with ridges or corrugations).
Composite material (rigidified laminate or xe2x80x9csandwichxe2x80x9d type).
Frontal area
Mainly metallic material (Al alloys or CSP-DB (super-plastic shaping and bonding diffusion)).
Composite materials (new materials with a high degree of impact energy absorption).
Until a few decades ago, the leading edges of almost all the supporting surfaces of aircraft had a metallic structure with a configuration of an outer covering of plate and an inner structure of the multi-rib type.
With the arrival of composite materials this type of structure has continued to be applied in a carbonised version but, at the same time, a step has been taken towards new solutions with structures of the xe2x80x9csandwichxe2x80x9d type (honeycomb), with a new configuration in which there are practically no ribs.
Up to the present time, the applicant has applied a very significant number of these solutions to different types of aeroplanes. Among them, many have been with carbon fibre, as in theory, the good rigidity and low weight of this material makes it extremely attractive. However, against this, there is the fragility of this material and the need to withstand the impact of birds, which affects the weight, giving a final result of doubtful effectiveness. The applicant""s experience has been developed and confirmed with many trials in a large-scale technological project called GSS (Large Supporting Surfaces). This work has made the following trinomial compatible in a reasonable and effective way:
Carbon fibre (less weight)
Minimal number of ribs (lower cost)
Large impact capacity (need for safety).
Based on his own previous technology, the applicant has developed the present invention for obtaining leading edges of supporting surfaces in aircraft, in which two basic concepts (primary structure and frontal area of the leading edge) are simultaneously applied, with very specific purposes and configurations.
Specifically, the invention has developed a leading edge of aircraft supporting surfaces in which the primary structure comprises an aerodynamic outer part made up of two covering panels, one upper and one lower, and an inner part made up of two spars, at least one of which has a diagonal configuration and another which closes in its frontal part, the two covering panels and the two spars constituting the strong and rigid structural combination of the leading edge. The frontal area of the leading edge is made up of one or two elemental parts which determine a kind of roof tile which closes the outer aerodynamic contour of the leading edge profile and which provides the latter with good behaviour against erosion, operational impacts on the ground and, above all, impact of birds in flight.
The covering panels, the spars and the frontal area of the leading edge may be made of a material selected from metals and composite materials.
Preferably, the composite material of said panels is selected from composite materials of the xe2x80x9csandwichxe2x80x9d type and solid laminates with stiffeners, obtained in their different forms, such as fabrics, tapes, RTM, RFI or thermoplastics, or alternatively, the metal of the covering panels consists of metallic plate with stiffeners or mechanised portions.
Moreover, the metal of the spars may consist of rigidified metallic plate with ridges or corrugations, and the composite material of said spars is selected from composite materials of the xe2x80x9csandwichxe2x80x9d type and rigidified laminates.
With respect to the frontal area of the leading edge, the metal of the same may be chosen from among aluminium alloys and titanium elements manufactured by a CSP-DB process, and the composite materials of said central area may be selected from new materials with a high degree of impact energy absorption.
Preferably, the covering panels and the spars consist of a xe2x80x9csandwichxe2x80x9d type composite material made up of outer laminates of carbon fibre which have a core of the honeycomb type solidly joined to their inside. Alternatively, the covering panels and the spars may consist of a solid laminate constituting an outer covering and stiffeners arranged in the direction of the chord to stabilise the covering and give it rigidity.
According to an aspect of the invention, the inner diagonal and frontal spars may consist of metallic plate with dimples and voids for lightness.
According to another aspect of the invention, the lower covering panel of the leading edge may be fixed with screws so that it can be detached, in which case the diagonal spar is provided with voids to facilitate access and inspection of the inner enclosure of the leading edge once the lower covering panel has been detached, said diagonal spar being also fixed with rivets arranged in intermediate local positions.
According to a further aspect of the invention, the frontal area of the leading edge may consist of a single metallic plate which is chemically shaped and milled in local areas.
Alternatively, the frontal area of the leading edge may consist of two metallic plates of aluminium joined together, one inner and the other outer, configured according to the aerodynamic profile of the leading edge.
As another alternative, the frontal area may consist of a titanium element manufactured by a CSP-DB process. In this case, the frontal spar of the leading edge may form part of the titanium element of the frontal area.